Currently, large parts of the panels that form the wings, stabilizers and parts of the fuselage of an airplane are fabricated of composite material, due to the reduced weight and which doubtlessly provide advantages in the field of aeronautics.
These panels are able to suffer, in the assembly process (which includes the handling, their own assembly, and the transfers) accidental damage. Following these said damages, the piece or panel is repaired or rejected. Particularly in the case of airplane panels with large dimensions, the probability of damage occurring to the piece is much higher than in other types of pieces, being the damaged area of a clearly irregular contour, upon being these irregular contours the area most likely to suffer these damages, the said damages are repaired through a very laborious process, elevating the cost and time to carry out the said process of repair.
Presently, the procedures of repair of the composite material aircraft panels are very expensive and laborious.
Thus, in the case of repairing large composite airplane panels, the following process at present consists of primarily cleaning the affected or damaged area, sanding the damaged area and step by step drawing out the fiber fabric in an area much bigger than the initially damaged area. The fabric which is removed will be returned to the same area and, with respect to the composite material, it should be placed again in the panel and processed and then cured in an autoclave. This entire process involves a large amount of time and expense, and there are not always available the necessary materials and infrastructures. This type of repair requires a specific technology, after which the area affected by the repair is much bigger than the initial damaged area.
The present invention offers a solution to the aforementioned problems.